The present invention is directed to hot-section turbine components in gas turbine engines, and more specifically to cooling structures of turbine components.
A gas turbine engine includes, in sequential flow order: a compressor, a combustor, and one or more turbines. During operation, air is compressed through the compressor and is mixed with fuel in the combustor. The fuel is ignited in the combustor, generating hot, high energy combustion gases which flow downstream through the turbine stages. The turbine stages extract energy from these combustion gases.
To prolong the service life of turbine blades and other hot section components and to reduce engine operating cost, portions of these components, for example turbine blade tips, often employ “active cooling”. This type of cooling is effected by bleeding off pressurized air at a relatively low temperature from some other portion of the engine such as the compressor, and then discharging the bleed air through cooling holes to form a protective film.
One problem with active cooling is that the use of bleed air is expensive in terms of overall fuel consumption.
Another problem with active cooling, particularly for turbine blade tips, is that the cooling holes can be damaged during a “rub” event in which the blade tips contact the surrounding shroud, thus lowering cooling effectiveness.